The present invention relates to thermal barrier coating methods.
Aircraft turbine engine components such as superalloy turbine vanes and blades can be protected against extremely high temperatures produced in aircraft turbine combustion chambers. Thermal barriers of protective nonoxide ceramic coatings are often applied to the metallic turbine engine components. Such thermal barriers may comprise refractory ceramic materials of silicon carbide and silicon nitride which can have high fracture strength in high temperature environments, have good thermal shock resistance, and can be fabricated into required shapes and sizes. See for example the following U.S. Pat. No. 5,626,923 to Fitzgibbons et al.; U.S. Pat. No. 5,639,531 to Chen et al.; U.S. Pat. No. 4,553,455 to Craig et al.; U.S. Pat. No. 4,405,660 to Ulion et al. and U.S. Pat. No. 5,514,482 to Strangrnan, all incorporated by reference herein.